(1) Field of the Invention
The present invention relates to a turbine engine component having an integrated system for cooling the platform, the tip, and the main body of an airfoil portion of the component.
(2) Prior Art
FIG. 1 depicts an engine arrangement 10 illustrating the relative location of a high pressure turbine blade 12. FIGS. 2 and 3 depict the main design characteristics of a typical conventionally cooled high-pressure blade 12. In general, cooling flow passes through these blades by means of internal cooling channels 14 that are turbulated with trip strips 16 for enhancing heat transfer inside the blade. The cooling effectiveness of these blades is around 0.50 with a convective efficiency of around 0.40. It should be noted that cooling effectiveness is a dimensionless ratio of metal temperature ranging from zero to unity as the minimum and maximum values. The convective efficiency is also a dimensionless ratio and denotes the ability for heat pick-up by the coolant, with zero and unity denoting no heat pick-up and maximum heat pick-up respectively. The higher these two dimensionless parameters become, the lower the parasitic coolant flow required to cool the high-pressure blade. In other words, if the relative gas peak temperature increases from 2500 degrees Fahrenheit to 2850 degrees Fahrenheit, the blade cooling flow should not increase and if possible, even decrease for turbine efficiency improvements. That objective is extremely difficult to achieve with current cooling technology which is shown schematically in FIGS. 2 and 3. In general, for such an increase in gas temperature, the cooling flow would have to increase more than 5% of the engine core flow. The metal temperature in the embodiment of FIG. 3 is about 2180 degrees Fahrenheit. This level of temperature is considered above the target limit.